A type of prior art gas turbine engine has a compressor, a fuel source, a combustion air source, a casing, and a combustor to prepare a heated fluid from fuel and combustion air. The combustor is connected to the fuel source, to the combustion air source and also to the compressor, which changes the pressure of the heated fluid before feeding it to the turbine. Practically the entire fluid flow from the compressor is directed to the combustor. The engine has a turbine rotor disk with blades that receive the heated fluid from the combustor. As the turbine rotor disk rotates during engine operation, the heated fluid flow coming from the combustor has to be directed at an angle to the blades to ensure smooth entry conditions. This is done using stator vanes that are positioned at a certain angle and direct the heated fluid from the combustor to the turbine rotor disk in a manner compatible with rotor disk rotation. This gas turbine engine is disclosed in U.S. Pat. No. 3,826,084 to Branstrom et al.
The stator vane angle normally is chosen to optimize efficiency based on prevailing turbine rotor disk operating conditions (speed). This solution is quite acceptable for gas turbine engines that have more or less steady operating conditions, such as when used for power generation. In applications where the load upon the gas turbine engine is steady, the turbine rotor disk rotates at a steady speed, and the entry angle for the blades remains unchanged thus minimizing losses. If, on the other hand, this gas turbine engine is used to power a vehicle, the situation is radically different. In that application, the turbine rotor disk speed will vary within a broad range depending on vehicle load. Consequently, the entry angle also varies within a broad range under load fluctuations, which leads to greater losses. This problem could not be solved by using the conventional approach with the stator vanes. It is possible to use controllable stator vanes to change the entry angle at the blades, but it is a very complicated and expensive solution given the high temperatures downstream of the combustor and space limitations. As a result, the gas turbine engine would have high losses in vehicle applications. Moreover, the stator and vanes occupy an additional space and make the engine design more complicated and expensive. The use of controllable vanes makes the engine less reliable.
The above problems are eliminated in our pending patent application Ser. No. 09/161,104 for a gas turbine engine, filed Sep. 25, 1998. A gas turbine engine disclosed in the above-mentioned as turbine engine has a device to admit a rotating fluid flow from an annular space in the casing to the inlet portion of a combustor to form a rotating fluid flow in the inlet portion of the combustor. The rotating fluid flow is formed in the annular space of the casing by supplying a fluid from a compressor to the blades of the turbine rotor disk. The fluid flow in an inlet portion of the combustor has a spin imparted to the fluid by rotation of the turbine rotor blades.
The disadvantage of the above gas turbine engine is an unstable clearance between the turbine rotor blade and the combustor. This is due largely to temperature fluctuations, temperature induced changes in rotor disk diameter, axial rotor disk displacement and wear on the bearings. These clearance changes between the rotor blades and combustor result in fluctuations of the overall fluid flow, including the flow through the combustor. It is imperative to use very stringent manufacturing tolerances in order to compensate for these negative phenomena to the maximum extent possible. In addition, the spin imparted to the fluid in the combustor by the turbine blades can be too strong if the blade radius is large. If the fluid spin in the combustor is too strong, the combustor may suffer high hydraulic losses, which lowers efficiency.
The problems indicated above are solved in the gas turbine engine of this invention.